XFOIL Version 6.94 Calculated polar for: EPPLER 540 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.2719 0.01357 0.00843 -0.0548 0.7354 0.7675 0.500 0.3251 0.01344 0.00827 -0.0550 0.7295 0.7693 1.000 0.3754 0.01325 0.00812 -0.0545 0.7227 0.7713 1.500 0.4301 0.01305 0.00790 -0.0549 0.7157 0.7729 2.000 0.4858 0.01295 0.00778 -0.0557 0.7084 0.7746 2.500 0.5358 0.01275 0.00765 -0.0552 0.6996 0.7772 3.000 0.5925 0.01252 0.00740 -0.0561 0.6909 0.7791 3.500 0.6437 0.01232 0.00727 -0.0560 0.6808 0.7812 4.000 0.6987 0.01208 0.00704 -0.0566 0.6698 0.7832 4.500 0.7487 0.01191 0.00696 -0.0561 0.6567 0.7850 5.000 0.8000 0.01175 0.00683 -0.0558 0.6419 0.7870 5.500 0.8448 0.01147 0.00663 -0.0540 0.6225 0.7888 6.000 0.8837 0.01133 0.00660 -0.0509 0.5961 0.7905 6.500 0.9145 0.01135 0.00662 -0.0462 0.5557 0.7925 7.000 0.9201 0.01172 0.00677 -0.0366 0.4913 0.7950 7.500 0.9151 0.01281 0.00754 -0.0257 0.4156 0.7978 8.000 0.9093 0.01427 0.00868 -0.0157 0.3451 0.8008 8.500 0.9093 0.01590 0.01006 -0.0076 0.2821 0.8040 9.000 0.9149 0.01767 0.01159 -0.0012 0.2260 0.8061 9.500 0.9222 0.01950 0.01323 0.0045 0.1757 0.8087 11.000 0.9531 0.02593 0.01923 0.0177 0.0644 0.8169 11.500 0.9655 0.02837 0.02158 0.0211 0.0437 0.8198 12.000 0.9797 0.03086 0.02407 0.0239 0.0320 0.8227 12.500 0.9952 0.03345 0.02671 0.0262 0.0247 0.8255 14.000 1.0244 0.04338 0.03702 0.0323 0.0099 0.8360 14.500 1.0310 0.04749 0.04133 0.0336 0.0079 0.8397 15.000 1.0299 0.05265 0.04672 0.0345 0.0069 0.8435 16.000 1.0283 0.06404 0.05852 0.0340 0.0059 0.8516 16.500 1.0189 0.07138 0.06613 0.0329 0.0055 0.8561 17.000 1.0150 0.07860 0.07365 0.0309 0.0053 0.8612 17.500 1.0084 0.08672 0.08206 0.0280 0.0051 0.8661 18.000 0.9985 0.09565 0.09132 0.0244 0.0050 0.8721 18.500 0.9854 0.10561 0.10161 0.0198 0.0048 0.8796 19.500 0.9479 0.12848 0.12523 0.0079 0.0048 0.9068 20.000 0.9268 0.14252 0.13973 -0.0017 0.0049 1.0000 20.500 0.8980 0.15810 0.15559 -0.0113 0.0050 1.0000 21.000 0.8600 0.17690 0.17467 -0.0223 0.0052 1.0000 23.000 0.7023 0.27184 0.27003 -0.0941 0.0111 0.8368 23.500 0.7131 0.27960 0.27783 -0.0984 0.0102 0.8389 24.000 0.7241 0.28761 0.28587 -0.1024 0.0097 0.8413 24.500 0.7315 0.29850 0.29678 -0.1075 0.0085 0.8430 25.000 0.7423 0.30624 0.30455 -0.1116 0.0078 0.8453 25.500 0.7526 0.31511 0.31345 -0.1157 0.0076 0.8475 26.000 0.7592 0.32676 0.32513 -0.1205 0.0066 0.8498 26.500 0.7675 0.33542 0.33384 -0.1244 0.0061 0.8528 27.500 0.7840 0.35309 0.35158 -0.1321 0.0056 0.8592 28.000 0.7913 0.36437 0.36289 -0.1365 0.0053 0.8624 28.500 0.7993 0.37434 0.37288 -0.1406 0.0049 0.8660 29.000 0.8060 0.38349 0.38208 -0.1443 0.0045 0.8704 29.500 0.8118 0.39220 0.39083 -0.1478 0.0043 0.8753 30.000 0.8179 0.40081 0.39949 -0.1514 0.0041 0.8805 30.500 0.8246 0.40959 0.40830 -0.1551 0.0039 0.8860 31.000 0.8283 0.41823 0.41698 -0.1580 0.0039 0.8927 31.500 0.8314 0.42725 0.42604 -0.1610 0.0039 0.9005 32.000 0.8349 0.43645 0.43527 -0.1641 0.0038 0.9091 32.500 0.8333 0.44184 0.44069 -0.1650 0.0037 0.9252